The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases flow downstream through turbine stages which extract energy therefrom for powering the compressor and producing useful work, such as powering a fan for propelling an aircraft in flight.
A high pressure turbine is disposed immediately downstream from the combustor and receives the hottest combustion gases therefrom. The first stage turbine rotor blades have hollow airfoils which are supplied with a portion of air bled from the compressor for use as a coolant in removing heat from the blades during operation.
Each airfoil includes pressure and suction sidewalls joined together at opposite leading and trailing edges, and extending from root to tip. A platform is disposed at the airfoil root and defines a portion of the radially inner flowpath for the combustion gases. And, a dovetail is integrally joined to the platform for mounting the individual blades in corresponding dovetail slots in the perimeter of a rotor disk.
Since the airfoil leading edge first engages the hot combustion gases, it requires substantial cooling for obtaining a useful blade life. Heat load from the combustion gases varies around the outer surface of the airfoil from the leading to trailing edges, and along the pressure and suction sidewalls. Various cooling circuits are provided inside the airfoil for cooling the different portions thereof. The different portions of the airfoil therefore operate at different temperatures, which introduces thermal stress therein that affects low cycle fatigue life of the blade.
Airfoil cooling may be effected using convection cooling, film cooling, or impingement cooling, or combinations thereof. The leading edge of a first stage turbine airfoil typically includes several rows or columns of film cooling holes fed by a common leading edge flow chamber or channel. Other film cooling holes and trailing edge holes may be fed by corresponding internal channels, such as multi-pass serpentine cooling channels.
In one conventional configuration, the leading edge chamber may be fed by a single flow channel located therebehind and separated therefrom by an intervening cold rib or bridge. The cold bridge includes a row of impingement holes which direct the air coolant in impingement against the backside of the leading edge for enhanced cooling thereof. However, the air flowing through the impingement supply channel is heated as it flows therethrough which correspondingly reduces impingement cooling effectiveness at the leading edge.
Since the cold bridge is disposed inside the airfoil and is itself cooled by the coolant therein, its temperature is substantially lower than that of the sidewalls of the airfoil around the leading edge. Accordingly, significant differential thermal expansion is effected between the sidewalls and the cold bridge which in turn creates large thermal stresses which adversely affect fatigue life.
The airfoil may include additional film cooling holes disposed in either sidewall downstream of the leading edge, which are typically referred to as gill holes. Since the gill holes are typically provided with a common source of coolant inside the airfoil, and the pressure of the combustion gases outside of the airfoil varies, backflow margin across the gill holes may vary on opposite sides of the airfoil.
Backflow margin is defined as the pressure of the coolant inside the airfoil divided by the local pressure of the combustion gases outside the airfoil as experienced by each of the gill holes. Sufficient backflow margin must be maintained to prevent ingestion of the hot combustion gases into the airfoil, and ensure continuous discharge of the coolant through the gill holes.
Since the minimum required backflow margin must be set at the airfoil leading edge, the backflow margin of the lower pressure suction sidewall gill holes may be undesirably high.
Accordingly, it is desired to provide a gas turbine engine turbine blade having improved leading edge cooling which addresses one or more of these typical design problems.
A turbine blade includes an airfoil having a leading edge flow chamber disposed behind a leading edge in front of a pair of side channels. The leading edge channel receives coolant from the pressure side channel in isolation from the suction side channel.